Gas turbine combustion system and method of flame stabilization in such a system

ABSTRACT

A gas turbine combustion system is provided. In an embodiment, the system includes a first radial inflow swirler having first radial outer intake openings, first radial inner outlet openings and first flow passages, each first flow passage including a first angle (a) with respect to the radial direction, a second radial inflow swirler having second radial outer intake openings, each second flow passage including a second angle (p) with respect to the radial direction, where the radial outer circumference of the second radial inflow swirler has a diameter that is smaller than the diameter of the radial inner circumference of the first radial inflow swirler and the second radial inflow swirler is located coaxially with and radially inside the first radial inflow swirler. The first angle (a) has a different sign than the second angle (p) with respect to the radial direction.

CROSS REFERENCE TO RELATED APPLICATIONS

This application is the US National Stage of International ApplicationNo. PCT/EP2012/074412 filed Dec. 5, 2012, and claims the benefitthereof. The International Application claims the benefit of EuropeanApplication No. EP 12159203 filed Mar 13, 2012. All of the applicationsare incorporated by reference herein in their entirety.

FIELD OF INVENTION

The present invention relates to a gas turbine combustion system and toflame stabilisation in a gas turbine combustion system. In particular,the invention relates to flame stabilisation in swirl stabilizeddiffusion flames.

BACKGROUND OF INVENTION

Although conventional diffusion flames that are swirl stabilised are notas prone to flame instabilities as are the flames in dry low emissionburners (DLE-burners), in which the air/fuel ratio is at or nearstoichiometric in order to reduce pollutants, the conventional burnersstill need a proper stable mixing to avoid any flameouts. In particular,if conventional burners are to be driven with a fuel containing H₂,which is for example present to a considerable amount in syngas or cokeoven gas (COG), flame stabilisation is still an issue because thesegases will lead to higher flame speeds which might end up in moreflameouts.

Multiple swirler concepts for manipulating mixing of fuel and air in gasturbine combustion systems are known from the state of the art. Forexample Bassam Mohammad and San-MouJeng “The Effect of Geometry on theAerodynamics of a Prototype Gas Turbine Combustor”, Proceedings of ASMETurbo Expo 2010: Power for Land, Sea and Air GT 2010, Jun. 14-18, 2010,Glasgow, UK, EP 2 192 347 A1 and U.S. Pat. No. 6,253,555 B1 describecombustion systems in which two radial inflow swirlers are arrangedaxially along a combustor central axis. In these combustion systems eachradial swirler is used by different airstreams. While in the first twomentioned documents both swirlers produce a swirl with the samerotational direction the swirlers of U.S. Pat. No. 6,253,555 B1 produceswirls of different rotational direction.

Yehida A. Eldrainy, et al. “A Multiple Inlet Swirler for Gas TurbineCombustors”, World Academy of Science, Engineering and Technology, 53,2009 describe a combustion system, in which a radial inflow swirler andan axial inflow swirler are combined.

U.S. Pat. No. 6,311,496 B1 describes a gas turbine combustion systemwith two radial inflow swirlers that are successively used by theairstream.

US 2005/00257530 A1 shows a fuel-air mixing apparatus in which tworadial inflow swirlers are used which have different radii and which aredisplaced relative to each other in an axial direction. The firstswirler is located upstream from plain jet orifices of a fuel deliveryline in said axial direction whereas the second swirler is locateddownstream from the plain jet orifices in said axial direction.

EP 0 939 275 A2 describes a fuel nozzle and nozzle guide for a gasturbine engine. The fuel nozzle includes a radial inflow swirler and anannular air passage leading from the swirler to a combustion chamber ofthe gas turbine engine. In addition, the fuel nozzle includes an axialswirler and a tubular air passage that leads to the combustion chamberand is encircled by the annular air passage. The radial swirler and theaxial swirler may produce co-swirls or counter-swirls. The nozzle guideincludes a radial inflow swirler and a frustroconical air passageleading from the swirler to the combustion chamber. The radial inflowswirler and the frustroconical air passage of the nozzle guide arecoaxial to the radial inflow swirler and the annular air passage of thenozzle. The radial inflow swirler of the nozzle guide may provide aco-swirl or a counter-swirl relative the swirl from the swirlers of thefuel nozzle.

EP 0 660 038 A2 shows a mixing duct with two annular arrays of swirlervanes which are separated by an annular divider. They may produceco-swirls or counter-swirls.

However, in particular for combustion systems using fuel gas withhydrogen H₂ like syngas or coke oven gas there is still need ofimproving flame stabilisation.

SUMMARY OF INVENTION

Hence, it is an objective of embodiments of the present invention toprovide a design for a gas turbine combustion system with increasedstability of diffusion flames. It is a further objective of embodimentsof the present invention to provide a method of flame stabilisation in agas turbine combustion system, in particular for diffusion flames.

The first objective is achieved by a gas turbine combustion system asclaimed in an independent claim. The second objective is achieved by amethod of flame stabilisation in a gas turbine combustion system asclaimed in another independent claim. The depending claims containfurther developments of the invention.

An inventive gas turbine combustion system comprises a central axis anda radial direction with respect to said central axis, a first radialinflow swirler and a second radial inflow swirler.

The first radial inflow swirler has radial outer intake openings locatedat a radial outer circumference of the first radial inflow swirler. Theradial outer intake openings of the first radial inflow swirler arerefered to as first radial outer intake openings in the following.Moreover, the first radial inflow swirler has outlet openings located ata radial inner circumference of the first radial inflow swirler. Theseoutlet openings are referred to as first radial inner outlet openings inthe following. Flow passages, named first flow passages in thefollowing, extend from the first radial outer intake openings to thefirst radial inner outlet openings. Each first flow passage includes afirst angle with respect to the radial direction.

The gas turbine combustion system further comprises a second radialinflow swirler having radial outer intake openings which are located ata radial outer circumference of the second radial inflow swirler andwhich are referred to as second radial outer intake openings in thefollowing. In addition, the second radial inflow swirler has radialinner outlet openings, which are referred to as second radial inneroutlet openings in the following and which are located at a radial innercircumference of the second radial inflow swirler. Flow passages, namedsecond flow passages in the following, extend from the second radialouter intake openings to the second radial inner outlet openings. Eachsecond flow passage includes an angle with respect to the radialdirection. This angle is referred to as a second angle in the following.In a particular embodiment of the inventive gas turbine combustionsystem, the number of second flow passages may be identical to thenumber of first flow passages.

The radial outer circumference of the second radial inflow swirler has adiameter that is at least slightly smaller than the diameter of theradial inner circumference of the first radial inflow swirler, and thesecond radial inflow swirler is located coaxially with and radiallyinside an opening formed by the inner circumference of the first radialinflow swirler, i.e. inside a space encircled by the radial innercircumference of the first radial inflow swirler.

According to an embodiment of the invention, the first angle has adifferent sign than the second angle with respect to the radialdirection. In other words, the second radial inflow swirler produces aswirl counter rotating with respect to the swirl generated by the firstradial inflow swirler. The counter rotation produced by the two swirlersleads to a more uniform mixing of an oxidant, like in particular theoxygen in the air, and fuel and to a stable flame which has theadvantages of lesser flameouts, a more distributed mixing of fuel andthe oxidant, a better control of the combustion burner, lesser hotspotsand a lower heat load across the metal surfaces like, for example, thecombustor walls. In a further development of the inventive gas turbinecombustion system, the first angle and the second angle may have thesame absolute value so that they only differ in their orientation withrespect to the radial direction.

Preferably, fuel injection openings are located in the second radialinflow swirler and are open towards the second flow passages. Morepreferably, the fuel injection openings are located inside the secondflow passages, in particular in the radial outer half of the second flowpassages, preferably in the outer third of the second flow passages. Byinjecting fuel into the second flow passages a particular effectiveflame stabilisation can be achieved.

In an advantageous further development of an embodiment of the inventivecombustion system, a radial gap may be present between the radial innercircumference of the first radial inflow swirler and the radial outercircumference of the second radial inflow swirler. In this case, theflow cross section of the second flow passages may be smaller than theflow cross section of the first flow passages since part of the fluidcan be introduced into a combustion chamber through the radial gap whileanother part will be introduced into the combustion chamber through thesecond radial inflow swirler.

According to a second aspect of the present invention, a method of flamestabilisation in a gas turbine combustion system is provided. In thecombustion system, a fluid flows along a flow path with a radialcomponent from a fluid inlet to a fluid outlet. The fluid is a fluidthat comprises an oxidant, and a fuel is mixed with the fluid thatcomprises an oxidant so as to transform the fluid into a mixturecomprising fuel and the oxidant. When air is used as the fluid (thatcomprises oxygen as the oxidant) the fluid is transformed into afuel/air mixture. A first swirl with a first rotational direction isintroduced into the flowing fluid in a radial upstream section of theflow path by passing the fluid through the first radial inflow swirlerof the gas turbine combustion system to generate a swirling fluid.Moreover, in a radial downstream section of the flow path a second swirlwith a second rotational direction is introduced into at least a portionof the fluid that exits the outlet openings of the first radial inflowswirler by passing the fluid through the second radial inflow swirler ofthe gas turbine combustion system to generate a swirling fluid.According to an embodiment of the inventive method, the secondrotational direction represents a counter rotation with respect to thefirst rotational direction. By introducing a counter rotation a betterstability of the diffusion flame and a more uniform mixing of fuel andthe oxidant can be achieved, as mentioned above with respect to anembodiment of the inventive combustion system. This is, in particular,true if the fuel contains hydrogen.

Embodiments of the inventive method are particularly effective inimproving flame stability and uniform mixing of fuel and oxidant if fuelis introduced into the fluid where the second swirl is generated. Inparticular, the fuel is introduced into the fluid at a location wheregeneration of the second swirl begins.

According to a further embodiment of the invention, no second swirl isintroduced into a portion of the fluid.

BRIEF DESCRIPTION OF THE DRAWINGS

Further features, properties and advantages of embodiments of thepresent invention will become clear from the following description ofembodiments in conjunction with the accompanying drawings.

FIG. 1 schematically shows a combustor arrangement for a gas turbinewith an inventive combustion system and a combustion chamber.

FIG. 2 shows the combustion system as seen from the combustion chamber.

DETAILED DESCRIPTION OF INVENTION

Embodiments of an inventive combustion system will be described withrespect to FIGS. 1 and 2 in the context of a combustor arrangementincluding an inventive combustion system. The inventive combustionsystem is adapted for performing the inventive method of flamestabilisation in a gas turbine combustion system which will also bedescribed with respect to FIGS. 1 and 2.

FIG. 1 shows part of a combustor arrangement in a sectional view. Thecombustor arrangement comprises a combustion chamber 3 and a combustionsystem 1 that is connected to a combustion chamber 3 via a smallpre-chamber 5. The pre-chamber is sometimes also called transitionsection and may be part of the combustion system 1 like in the presentembodiment. However, the pre-chamber 5 may as well be a part of thecombustion chamber 3 or a distinct part that is neither part of thecombustion system 1 nor of the combustion chamber 3.

The combustion system 1 comprises a first radial inflow swirler 7 that,shows rotational symmetry with respect to a central combustor axis A.The first radial inflow swirler is equipped with a number of vanes 9that are distributed along the circumferential direction of the swirler7 and are spaced apart from each other. Flow passages 11 are formedbetween neighbouring vanes 9. Each flow passage 11 extends from a firstradial outer intake opening 13 located at a radial outer circumferenceof the swirler 7 to a first radial inner outlet opening 15 located at aradial inner circumference of the swirler 7. The flow passages 11 of thefirst swirler 7 are angled with respect to the radial direction of theswirler with a first angle a so that a swirl is imparted to a fluidflowing through the flow channel 11.

The combustion system 1 further comprises a second radial inflow swirler17 that, like the first radial inflow swirler, shows radial symmetry.However, the second radial inflow swirler 17 has an outer circumferencethe diameter of which is smaller than the inner circumference of thefirst radial inflow swirler 11. The second radial inflow swirler 17 islocated inside an opening formed by the inner circumference of the firstradial inflow swirler 7 so that a fluid that exits the outlet openings15 of the first radial inflow swirler 7 is directed towards the secondradial inflow swirler 17.

Like the first radial inflow swirler 7, the second radial inflow swirler17 comprises a number of vanes 19 that are distributed incircumferential direction of the swirler such that second flow passages21 are formed between them. Each second flow passage 21, i.e. each flowpassage of the second radial inflow swirler 17, extends from a secondradial outer intake opening 23 located at the radial outer circumferenceof the second swirler to a second radial inner outlet opening 25, i.e.an outlet opening of the second swirler 17 that is located at the innercircumference of the second radial inflow swirler 17. The radial outerintake openings 23 of the second radial inflow swirler 17 show towardthe radial inner circumference of the first radial inflow swirler 7, inwhich the radial inner outlet openings 15 of the first radial inflowswirler 7 are located. Hence, a fluid exiting the first radial inflowswirler 7 can enter the second radial inflow swirler 17.

The flow channels 21 of the second radial inflow swirler 17 include anangle with the radial direction (denominated β in FIG. 2) which has, inthe present embodiment, the same absolute value as the angle of the flowchannels 11 of the first radial inflow swirler 7 but a different sign.Hence, the flow channels 11 of the first radial inflow swirler 7 imparta clockwise swirl to a flowing fluid and the flow channels 21 of thesecond radial inflow swirler 17 impart a counter-clockwise swirl to afluid flowing therethrough, or vice versa.

Both swirlers 7, 17 are mounted to a base plate 31 such that they arearranged coaxially with each other and with respect to the combustoraxis A. Hence, in radial direction the second radial inflow swirler 17is surrounded by the first radial inflow swirler 7. Moreover, in thepresent embodiment the radial inflow swirlers 7, 17 are arranged suchthat a radial gap 27 is formed between the inner circumference of thefirst radial inflow swirler 7 and the outer circumference of the secondradial inflow swirler 17.

Fuel channels 33 extend through the base plate 31 and lead to fuelopening 29 in the flow passages 21 of the second radial inflow swirler7. The fuel openings 29 are located in the outer half of the second flowpassages 21, preferably in the outer third of the second flow passages21, and more preferably in the outer fourth of the second flow passages21.

The first radial inflow swirler 7 is surrounded by a flow channel 35which allows feeding a fluid, in particular air or any other suitablefluid that comprises an oxidant, to the intake openings 13 of the firstradial inflow swirler.

During operation of a gas turbine air is fed to the intake openings 13of the first radial inflow swirler 7 through the flow channel 35. Theair then flows through the flow passages 11 of the first radial inflowswirler 7 whereby a first swirl (indicated by arrow 37) is imparted tothe flowing air.

Hence, in the present embodiment, the air swirls with a clockwiserotation after exciting the first swirler through the outlet openings15. A part of the clockwise swirling air reaches the pre-chamber 5through the radial gap 27. Another part of the clockwise swirling airenters the flow passages 21 of the second radial inflow swirler 17through the intake openings 23. Thereby, the intake openings 23 of thesecond radial inflow swirler generate turbulences in the flow channelsections adjoining the intake openings 15. The turbulences are generateddue to a reversal in rotation direction that is necessary for the air toenter the flow passages 21 of the second swirler 17. The turbulences arehighest in a flow passage zone adjoining the intake openings 23 of theflow passages.

A fuel gas like, for example, syngas or coke oven gas (COG) isintroduced into the turbulent airstreams in the second flow passages 21through the fuel holes 29. The strong turbulence leads to a highlyuniform mixing of fuel and air until the fuel/air mixture leaves thesecond flow channels 21 through the second outlet openings 25. Due tothe angle β the second flow passages 21 include with the radialdirection a second swirl (indicate by arrow 39) with a counter-clockwiserotation is imparted to the fuel/air mixture flowing through the secondflow passages 21.

A further effect of giving the angle of the flow channels of the firstand second swirlers a different sign with respect to the radialdirection is that the fuel/air mixture has a different direction ofrotation than the air entering the pre-chamber 5 through the gap 27 thatis present between both swirlers 7, 17 in the described embodiment. As aconsequence, the air rotating clockwise in the present embodiment canform an envelope around the fuel/air mixture rotating counterclockwisein the present embodiment which makes it more difficult for fuel/airmixture to reach the wall of the pre-chamber 5 and the combustionchamber 3, thereby reducing heat load across the metal surface of thecombustor wall. A further advantage is that turbulences are formed wherethe counter-clockwise swirling fuel/air mixture is in contact with theclockwise swirling air, which turbulences lead to a more distributedmixing of fuel and air. The mentioned effects contribute to leading toless flameouts and less hotspots, in particular with use of H₂containing gases like syngas or COG. In the end, this leads to a bettercontrollable combustion burner.

Aspects of the present invention has been described with respect to aspecific embodiment to describe a method of improve mixing of gas andair and to stabilise the flame by using the concepts of swirl strengthin diffusion flames to anchor it in a stabile way. In particular,counter rotating swirls are used to improve mixing and stabilising ofthe flame. However, the invention shall not be restricted to thespecific embodiment described with respect to the figures, sincedeviations therefrom are possible. For example, while in the Figuresboth swirlers have the same number of flow passages the second swirlercould have a higher or lower number of flow passages than the firstswirler. Moreover, the flow passages of both swirlers are angled by thesame absolute value with respect to the radial direction but with adifferent sign. In other embodiments it may be useful to also havedifferent absolute values of the angles between the flow passages andthe radial direction. A further possible deviation from the embodimentdescribed with respect to the figures is the number of fuel opening thatare present in each flow passage of the second swirler. While in thedescribed embodiment only one fuel opening is present in each flowpassage a higher number of fuel openings may be present as well.Moreover, the fuel openings do not need to be present in the base plate.Alternatively or additionally, fuel openings could be located in thesidewalls of the vanes. Since the location of the fuel openings isclosely related to the geometry of the swirler and the fuel to be usedthe exact position of the fuel openings may depend on the concretedesign of the first and second radial inflow swirler and/or on theintended use of the combustion system.

Since many deviations from the embodiment are possible, the presentinvention shall only be limited by the appended claims.

1. A gas turbine combustion system, comprising: a central axis (A) and aradial direction with respect to said central axis (A); a first radialinflow swirler comprising first radial outer intake openings located ata radial outer circumference of the first radial inflow swirler, thefirst radial inner outlet openings located at a radial innercircumference of the first radial inflow swirler, and first flowpassages extending from the first radial outer intake openings to thefirst radial inner outlet openings, each first flow passage including afirst angle (a) with respect to the radial direction; a second radialinflow swirler comprising second radial outer intake openings located ata radial outer circumference of the second radial inflow swirler, thesecond radial inner outlet openings located at a radial innercircumference of the second radial inflow swirler, and second flowpassages extending from the second radial outer intake openings to thesecond radial inner outlet openings, each second flow passage includinga second angle (β) with respect to the radial direction; wherein theradial outer circumference of the second radial inflow swirler hascomprises a diameter that is smaller than the diameter of the radialinner circumference of the first radial inflow swirler and wherein thesecond radial inflow swirler is located coaxially with and radiallyinside an opening formed by the inner circumference of the first radialinflow swirler so that a fluid that exits the outlet openings of thefirst radial inflow swirler is directed towards the second radial inflowswirler, and wherein the first angle (a) has a different sign than thesecond angle (β) with respect to the radial direction.
 2. The gasturbine combustion system as claimed in claim 1, wherein fuel injectionopenings are located in the second radial inflow swirler and are opentowards the second flow passages.
 3. The gas turbine combustion systemas claimed in claim 2, wherein the fuel injection openings are locatedinside the second flow passages.
 4. The gas turbine combustion system asclaimed in claim 3, wherein the fuel injection openings are located inthe radial outer half of the second flow passages.
 5. The gas turbinecombustion system as claimed in claim 1, wherein the number of secondflow passages is identical to the number of first flow passages.
 6. Thegas turbine combustion system as claimed in claim 1, wherein a radialgap is present between the radial inner circumference of the firstradial inflow swirler and the radial outer circumference of the secondradial inflow swirler.
 7. The gas turbine combustion system as claimedin claim 6, wherein the flow cross section of the second flow passagesis smaller than the flow cross section of the first flow passages. 8.The gas turbine combustion system as claimed in claim 1, wherein thefirst angle (a) and the second angle (β) have the same absolute value.9. A method of flame stabilisation in a gas turbine combustion system inwhich a fluid flows along a flow path with a radial component, by use ofa gas turbine combustion system as claimed in claim 1, wherein, thefluid is a fluid that comprises an oxidant and a fuel is mixed with thefluid that comprises an oxidant so as to transform the fluid into amixture comprising fuel and the oxidant; a first swirl with a firstrotational direction is generated in the flowing fluid in a radialupstream section of the flow path by passing the fluid through the firstradial inflow swirler of the gas turbine combustion system to generate aswirling fluid; in a radial downstream section of the flow path a secondswirl with a second rotational direction is generated in at least aportion of the fluid by passing said portion of the swirling fluidthrough the second radial inflow swirler of the gas turbine combustionsystem, wherein the second rotational direction represents a counterrotation with respect to the first rotational direction.
 10. The methodas claimed in claim 9, wherein fuel is introduced into the fluid wherethe second swirl is generated.
 11. The method as claimed in claim 10,wherein the fuel is introduced into the fluid at a location wheregeneration of the second swirl begins.
 12. The method as claimed inclaim 9, wherein no second swirl is introduced into a portion of thefluid.